Multi-cover gas turbine engine component

ABSTRACT

An airfoil for a gas turbine engine according to an example of the present disclosure includes, among other things, an airfoil body extending between leading and trailing edges in a chordwise direction and extending from a root section in a spanwise direction, and the airfoil body defining pressure and suction sides separated in a thickness direction. The airfoil body defines a recessed region extending inwardly from at least one of the pressure and suction sides, and the airfoil body includes one or more ribs that define a plurality of pockets within a perimeter of the recessed region. A plurality of cover skins is welded to the airfoil body along the one or more ribs to enclose respective ones of the plurality of pockets. The plurality of cover skins formed from a common cover having a perimeter that is dimensioned to mate with the perimeter of the recess. A method of forming a gas turbine engine component is also disclosed.

BACKGROUND

This disclosure relates to a gas turbine engine, and more particularlyto hollow gas turbine engine components.

Gas turbine engines can include a fan for propulsion air and to coolcomponents. The fan also delivers air into a core engine where it iscompressed. The compressed air is then delivered into a combustionsection, where it is mixed with fuel and ignited. The combustion gasexpands downstream over and drives turbine blades. Static vanes arepositioned adjacent to the turbine blades to control the flow of theproducts of combustion.

Some fans include hollow fan blades made of a metallic or compositematerial. Various techniques can be utilized to construct hollow fanblades, including attaching a cover to an airfoil body.

SUMMARY

An airfoil for a gas turbine engine according to an example of thepresent disclosure includes an airfoil body extending between leadingand trailing edges in a chordwise direction and extending from a rootsection in a spanwise direction, and the airfoil body defining pressureand suction sides separated in a thickness direction. The airfoil bodydefines a recessed region extending inwardly from at least one of thepressure and suction sides, and the airfoil body includes one or moreribs that define a plurality of pockets within a perimeter of therecessed region. A plurality of cover skins is welded to the airfoilbody along the one or more ribs to enclose respective ones of theplurality of pockets. The plurality of cover skins formed from a commoncover having a perimeter that is dimensioned to mate with the perimeterof the recess.

In a further embodiment of any of the foregoing embodiments, theplurality of cover skins are dimensioned to mate with a perimeter ofrespective ones of the plurality of pockets.

In a further embodiment of any of the foregoing embodiments, theplurality of cover skins includes a peripheral cover skin and one ormore localized cover skins. The peripheral cover skin includes theperimeter of the common cover such that the peripheral cover skinsurrounds the one or more localized cover skins in an installedposition.

In a further embodiment of any of the foregoing embodiments, theperipheral cover skin is welded to the airfoil body along the perimeterof the recessed region.

In a further embodiment of any of the foregoing embodiments, the one ormore localized cover skins are a plurality of localized cover skins thatare spaced apart from each other and from the perimeter of the recessedregion.

In a further embodiment of any of the foregoing embodiments, theplurality of localized cover skins includes a first cover skin. Thefirst cover skin has a plurality of branch segments extending from anelongated body.

In a further embodiment of any of the foregoing embodiments, the one ormore ribs include a plurality of ribs, each one of the plurality of ribsincluding a raised protrusion extending outwardly from a pedestalportion, the pedestal portion dimensioned to support an opposed pair ofthe plurality of cover skins, and the raised protrusion is dimensionedto extend between and space apart the opposed pair.

In a further embodiment of any of the foregoing embodiments, the airfoilis a fan blade.

A gas turbine engine according to an example of the present disclosureincludes a fan section that has a fan rotatable about an enginelongitudinal axis, a compressor section, a turbine section that drivesthe compressor section and the fan, and a plurality of airfoils eachincluding an airfoil body defining a recessed region extending inwardlyfrom a sidewall of the airfoil body. The sidewall includes a pluralityof ribs that divide the recessed region into a plurality of pockets. Aplurality of cover skins are formed from a common cover that isdimensioned with respect to an external surface contour of the airfoilbody. The plurality of cover skins have a peripheral cover skin and aplurality of localized cover skins mechanically attached to the airfoilbody along the plurality of ribs to enclose respective ones of theplurality of pockets, and the peripheral cover skin comprise a perimeterof the common cover.

In a further embodiment of any of the foregoing embodiments, the fancomprises the plurality of airfoils.

In a further embodiment of any of the foregoing embodiments, theperipheral cover skin is welded to a perimeter of the recessed regionsuch that the peripheral cover skin surrounds the plurality of localizedcover skins in an installed position, and the plurality of localizedcover skins are dimensioned to mate with a perimeter of respective onesof the plurality of pockets.

In a further embodiment of any of the foregoing embodiments, each one ofthe plurality of ribs includes a pedestal portion and a raisedprotrusion, and the raised protrusion is dimensioned to extend outwardlyfrom the pedestal portion to space apart an opposed pair of theplurality of localized cover skins in the installed position.

A method of forming a gas turbine engine according to an example of thepresent disclosure includes forming a recessed region in a sidewall of amain body, dividing the recessed region into a plurality of pocketsbetween one or more ribs such that the plurality of pockets aresurrounded by a perimeter of the recessed region, contouring a coveraccording to an external surface contour of the main body such that aperimeter of the cover is dimensioned to mate with the perimeter of therecessed region, dividing the cover to form a plurality of cover skinssubsequent to the contouring step, positioning the plurality of coverskins to enclose respective ones of the plurality of pockets, andwelding the plurality of cover skins to the main body along the one ormore ribs subsequent to the positioning step.

In a further embodiment of any of the foregoing embodiments, theplurality of cover skins includes a peripheral cover skin and one ormore localized cover skins. The peripheral cover skin includes aperimeter of the cover such that the peripheral cover skin surrounds theone or more localized cover skins subsequent to the positioning step.

In a further embodiment of any of the foregoing embodiments, the weldingstep includes welding the peripheral cover skin to the main body alongthe perimeter of the recessed region.

In a further embodiment of any of the foregoing embodiments, the one ormore ribs include a first set of ribs and a second set of ribs. Each oneof the second set of ribs extend from at least one of the first set ofribs, and the positioning step includes situating one or more of theplurality of cover skins over the second set of ribs.

In a further embodiment of any of the foregoing embodiments, the secondset of ribs are spaced apart from the plurality of cover skins.

In a further embodiment of any of the foregoing embodiments, each one ofthe one or more ribs includes a raised protrusion that extends outwardlyfrom a pedestal portion. The pedestal portion is dimensioned to supportan opposed pair of the plurality of cover skins, and the raisedprotrusion is dimensioned to abut against the opposed pair.

In a further embodiment of any of the foregoing embodiments, the raisedprotrusion is dimensioned to extend outwardly from external surfaces ofthe opposed pair subsequent to the positioning step, and the raisedprotrusion is consumed during the welding step.

In a further embodiment of any of the foregoing embodiments, theexternal surface contour of the main body and external surfaces of theplurality of cover skins cooperate to define a pressure side or asuction side of an airfoil.

The various features and advantages of this disclosure will becomeapparent to those skilled in the art from the following detaileddescription. The drawings that accompany the detailed description can bebriefly described as follows.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 illustrates an example turbine engine.

FIG. 2 illustrates a perspective view of a gas turbine engine componenthaving a plurality of cover skins.

FIG. 3 illustrates a section view of the component taken along line 3-3of FIG. 2 with the cover skins in an installed position.

FIG. 3A illustrates selected portions of the component of FIG. 3.

FIG. 3B illustrates a sectional view a stiffening rib of the componentof FIG. 3.

FIG. 4 illustrates the component of FIG. 2 with the cover skins removed.

FIG. 5 is a schematic view of airfoil span positions.

FIG. 6 is a schematic view of an airfoil section depicting a staggerangle at a span position of FIG. 5.

FIG. 7 illustrates a process for forming a gas turbine engine component.

FIG. 8 illustrates a gas turbine engine component including cover skinsformed from a common cover.

FIG. 8A illustrates a cover skin according to an example.

FIG. 8B illustrates a cover skin according to another example.

FIG. 9 illustrates one of the cover skins of FIG. 8 mounted on a supportstructure.

FIG. 9A illustrates a cover skin mounted on a support structure.

FIG. 10 illustrates another one of the cover skins of FIG. 8 mounted ona support structure.

FIG. 11 illustrates adjacent cover skins positioned relative to asupport rib of the component of FIG. 8.

FIG. 11A illustrates a support rib according to another example.

FIG. 11B illustrates a support rib according to yet another example.

FIG. 12 illustrates adjacent cover skins attached to the support rib ofFIG. 11.

FIG. 13 illustrates a cover skin attached to a main body of thecomponent of FIG. 8.

FIG. 14 illustrates a perspective view of a gas turbine engine componentaccording to another example.

FIG. 15 illustrates a perspective view of a gas turbine engine componentaccording to yet another example.

FIG. 16 illustrates a perspective view of a gas turbine engine componentaccording to another example.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. The fan section 22 drivesair along a bypass flow path B in a bypass duct defined within a nacelle15, and also drives air along a core flow path C for compression andcommunication into the combustor section 26 then expansion through theturbine section 28. Although depicted as a two-spool turbofan gasturbine engine in the disclosed non-limiting embodiment, it should beunderstood that the concepts described herein are not limited to usewith two-spool turbofans as the teachings may be applied to other typesof turbine engines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects, a first (or low) pressure compressor 44 and a first (orlow) pressure turbine 46. The inner shaft 40 is connected to the fan 42through a speed change mechanism, which in exemplary gas turbine engine20 is illustrated as a geared architecture 48 to drive a fan 42 at alower speed than the low speed spool 30. The high speed spool 32includes an outer shaft 50 that interconnects a second (or high)pressure compressor 52 and a second (or high) pressure turbine 54. Acombustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 may be arranged generallybetween the high pressure turbine 54 and the low pressure turbine 46.The mid-turbine frame 57 further supports bearing systems 38 in theturbine section 28. The inner shaft 40 and the outer shaft 50 areconcentric and rotate via bearing systems 38 about the engine centrallongitudinal axis A which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of the low pressure compressor, or aftof the combustor section 26 or even aft of turbine section 28, and fan42 may be positioned forward or aft of the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1 and less than about 5:1. Itshould be understood, however, that the above parameters are onlyexemplary of one embodiment of a geared architecture engine and that thepresent invention is applicable to other gas turbine engines includingdirect drive turbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and35,000 ft (10,668 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tram °R)/(518.7°R)]^(0.5). The “Lowcorrected fan tip speed” as disclosed herein according to onenon-limiting embodiment is less than about 1150 ft/second (350.5meters/second).

FIGS. 2-3 illustrate a gas turbine engine component 60 according to anexample. The component 60 can be incorporated in the engine 20 of FIG.1, for example. In the illustrated example of FIGS. 2-3, the component60 is an airfoil 61. The fan 42 of FIG. 1 can include a plurality ofairfoils or fan blades 43 rotatable about the engine longitudinal axisA, and the airfoil 61 can be one of the fan blades 43. Other types ofrotatable blades and airfoils, including struts and static vanes 45 inthe fan, compressor and turbine sections 22, 24, 28, mid-turbine frame57, and turbine exhaust case (TEC) 55 of FIG. 1 may benefit from theexamples disclosed herein which are not limited to the design shown.Other portions of the engine 20 including engine cases, generally planaror flat panels or structures, and other systems such as industrialturbines may benefit from the examples disclosed herein.

The airfoil 61 includes an airfoil section 62 extending in a spanwise orradial direction R from a root section 64 to a tip portion 66 (FIG. 2).The root section 64 is a shape that is configured to mount the airfoil61 in the engine, such as a dovetail shape. The tip portion 66 is aterminal end of the airfoil 61. Generally, one side of the airfoilsection 62 is a suction side SS and the other side is a pressure side PS(FIG. 3) separated in a thickness direction T. The pressure side PS hasa generally concave profile, and the suction side SS has a generallyconvex profile. The airfoil section 62 extends in the thicknessdirection T between the pressure and suction sides PS, SS to define anaerodynamic external surface contour of the airfoil section 62, asillustrated in FIG. 3. The airfoil 61 is rotatable about an axis ofrotation RR. The axis of rotation RR can be collinear with or parallelto the engine axis A of the engine 20.

The airfoil section 62 includes a main (or airfoil) body 68 that extendsin the radial direction R from the root section 64 to the tip portion66. The main body 68 extends in a chordwise direction X between aleading edge LE and a trailing edge TE. The main body 68 defines atleast one of the pressure and suction sides PS, SS. In the illustratedexample of FIGS. 2-3, the main body 68 defines both the pressure andsuction sides PS, SS.

The airfoil 61 includes a plurality of cover skins 70 disposed on asurface of the main body 68. The cover skins 70 are arranged to providea continuous surface with the suction side SS of the airfoil 61 when inan installed position, as illustrated by FIG. 3. In another example, thecover skins 70 are disposed on the pressure side PS of the airfoil 61.The airfoil 61 is illustrated with the cover skins 70 removed in FIG. 4for illustrative purposes. The component 60 includes two or more coverskins 70 along the pressure and/or suction sides PS, SS of the airfoilsection 62.

The main body 68 and cover skins 70 can be made out of metallicmaterials such as titanium. Other materials can be utilized, includingmetal alloys and metal matrix composites.

The cover skins 70 include a peripheral cover skin 72 and one or morelocalized cover skins 74. In the illustrated example of FIGS. 2-3, theairfoil 61 includes four localized cover skins 74 (indicated at 74-1 to74-4 in FIG. 2) distributed along the airfoil section 62. It should beunderstood that the component 60 can include fewer or more than fourlocalized cover skins 74 in accordance with the teachings disclosedherein, such as only one localized cover skin 74.

Referring to FIGS. 3-4 with continuing reference to FIG. 2, a sidewall76 of the main body 68 defines a recessed region 78 that is dimensionedto at least partially receive the cover skins 70. The recessed region 78extends inwardly from at least one of the pressure and suction sides PS,SS defined by the sidewall 76, such as the suction side SS asillustrated in FIGS. 3-4. The sidewall 76 includes one or more supportribs 80 that divide the recessed region 78 into, or otherwise define, aplurality of internal cavities or pockets 82 within a perimeter 78P ofthe recessed region 78. In the illustrated example of FIG. 4, thesidewall 76 defines at least five separate and distinct pockets 82(indicated at 82-1 to 82-5) bounded by the support ribs 80. The pockets82 can serve to reduce an overall weight of the component 60. Thesupport ribs 80 are dimensioned to abut against and support respectivecover skins 70.

The main body 68 can include a plurality of stiffening ribs 81 (FIG. 4)extending from the support ribs 80 and/or perimeter 78P of the recessedregion 78. In the illustrative example of FIG. 4, each stiffening rib 81interconnects one of the support ribs 80 with another support rib 80 orthe perimeter 78P of the recessed region 78. The stiffening ribs 81extend outwardly from a floor 83 (FIG. 4) of the respective pocket 82.The stiffening ribs 81 can serve to provide rigidity or stiffening tothe main body 68. The stiffening ribs 81 can be dimensioned to allow themain body 68 to flex to absorb impacts from foreign objection debris(FOD), which can reduce strain along weld joints.

The cover skins 72, 74 are dimensioned to mate with a perimeter ofrespective ones of the pockets 82 defined by the ribs 80 and/orperimeter 78P of the recessed region 78. The localized cover skins 74are dimensioned to enclose respective ones of pockets 82, and theperipheral cover skin 72 is dimensioned to enclose pocket 82-5 such thatthe peripheral cover skin 72 surrounds each one of the localized coverskins 74 in an installed position as illustrated by FIG. 3. The coverskins 70 are dimensioned such that each cover skin 72, 74 encloses onlyone of the pockets 82. In other examples, one or more of the cover skins72, 74 enclose two or more adjacent pockets 82. The localized coverskins 74 are spaced apart from each other and from the perimeter 78P ofthe recessed region 78 in the installed position, as illustrated in FIG.2.

The main body 68 defines a shelf 73 along the perimeter 78P of therecessed region 78, as illustrated by FIGS. 3A and 4. The shelf 73 isdimensioned to at least partially receive and mate with a perimeter ofthe peripheral cover skin 72.

Positioning the cover skins 70 can include situating one or more of thecover skins over the stiffening ribs 81, as illustrated by cover skin 70in FIG. 3B. Each stiffening rib 81 can be dimensioned to be spaced apartfrom adjacent ones of the cover skins 70 to define a clearance gap GG.

The cover skins 70 and pockets 82 can have various geometries, includinga generally elongated, oblong or racetrack shaped geometry asillustrated by localized cover skins 74-1 to 74-3 (FIG. 2) and pockets82-1 to 82-3 (FIG. 4). Other geometries can include a complex profile asillustrated by localized cover skin 74-4 (FIG. 2) and pocket 82-4. Inthe illustrated example of FIG. 2, localized cover skin 74-4 includes aplurality of branch segments 75 extending from an elongated body 77. Itshould be appreciated that each cover skin, rib and pocket and can havedifferent dimensions, shapes and/or orientations than illustrated in thedisclosed figures. The geometry of the pockets can be the same or candiffer. Other example geometries of the cover skins, ribs and pocketscan include circular or elliptical, rectangular and triangulargeometries. At least some of the ribs can be generally linear orcurvilinear.

The cover skins 70 are mechanically attached to the main body 68 alongthe support ribs 80 and/or perimeter 78P of the recessed region 78 toenclose respective ones of the pockets 82. Various techniques can beused to mechanically attach the cover skins 70 to the main body 68,including laser or electron beam welding, brazing, diffusion bonding orother fastening techniques. In the illustrated example of FIGS. 2-3 and3A, the peripheral cover skin 72 is welded to the main body 68 along theperimeter 78P of the recessed region 78 such that the peripheral coverskin 72 surrounds the localized cover skins 74 in the installedposition. The pockets 82 can substantially or completely free of anymaterial such that the airfoil section 62 is hollow in an assembledcondition.

Referring to FIG. 5, span positions of the airfoil section 62 areschematically illustrated from 0% to 100% in 10% increments to define aplurality of sections 67. Each section 67 at a given span position isprovided by a conical cut that corresponds to the shape of segments aflowpath (e.g., bypass flowpath B or core flow path C of FIG. 1), asshown by the large dashed lines. The airfoil section 62 extends from aplatform 69 (see also FIG. 1). In the case of an airfoil 61 with anintegral platform, the 0% span position corresponds to the radiallyinnermost location where the airfoil section 62 meets the fillet joiningthe airfoil section 62 to the platform 69. In the case of an airfoil 61without an integral platform, the 0% span position corresponds to theradially innermost location where the discrete platform 69 (see alsoFIG. 1) meets the exterior surface of the airfoil section 62. A 100%span position corresponds to a section of the airfoil section 62 at thetip portion 66.

Referring to FIG. 6, with continuing reference to FIG. 5, the airfoilsection 62 is sectioned at a radial position between the root section 64(FIG. 2) and tip portion 66. In examples, each airfoil section 62 isspecifically twisted about a spanwise axis in the radial direction Rwith a corresponding stagger angle α at each span position. Chord CD,which is a length between the leading and trailing edges LE, TE, formsstagger angle α relative to the chordwise direction X or a planeparallel to the axis or rotation RR. The stagger angle α can vary alongthe span of the airfoil section 62 to define a twist. For example, thetip portion 66 can define a stagger angle α relative to the root section64 that is greater than or equal to 5 degrees or 10 degrees, absolute.In some examples, the stagger angle α at the tip portion 66 relative tothe root section 64 is between 5-60 degrees, absolute, or more narrowlybetween 10-30 degrees, absolute, such that the airfoil section 62 istwisted about a spanwise axis as illustrated by the airfoil 61 of FIGS.2-3. The airfoil section 62 can be three-dimensionally twisted about thespanwise axis.

FIG. 7 illustrates a process of constructing or forming a gas turbineengine component in a flow chart 184. The process can be utilized toform a hollow component such as the airfoil 61 of FIGS. 2-4, or anothercomponent such as a solid airfoil, or another component of the engine 20including static vanes and struts, for example. Reference is made to thecomponent 60 of FIGS. 2-4 and a component 160 of FIG. 8 for illustrativepurposes. In this disclosure, like reference numerals designate likeelements where appropriate and reference numerals with the addition ofone-hundred or multiples thereof designate modified elements that areunderstood to incorporate the same features and benefits of thecorresponding original elements.

Referring to FIGS. 7-8, a main body 168 of the component 160 (shown indashed lines for illustrative purposes) can be prepared or otherwiseformed at step 184A. The main body 168 can be formed with respect to apredefined geometry, which can be defined with respect to one or moredesign criterion. Step 184A can include mounting the main body 168 to atool and machining internal and/or external surfaces of the main body168 with respect to the predefined geometry, such as the aerodynamicexternal surface contour CC of the airfoil section 62 of FIGS. 2-3characterized by a three-dimensional twist.

At step 184B, one or more surface features are formed or otherwisedefined in the sidewall 176 of the main body 168. In the illustratedexample of FIGS. 3-4, the surface features include the recessed region78, shelf 73, pockets 82 and ribs 80 distributed along the sidewall 76of the main body 68. The recessed region 78 is divided into two or morepockets 82 between one or more ribs 80 such that the pockets 82 aresurrounded by the perimeter 78P of the recessed region 78. The main body68 and surfaces features including the shelf 73, recessed region 78,ribs 80 and pockets 82 can be forged, cast, machined or produced byadditive manufacturing from a metal or metal alloy, for example.

A common cover 186 is formed at step 184C. The cover 186 can be forged,machined or produced by additive manufacturing from a metal or metalalloy, for example. In examples, the cover 186 is formed from a sheetmetal body having a substantially planar geometry. For the purposes ofthis disclosure, the term “substantially” means±3 percent of therespective value unless otherwise stated.

Forming the common cover 186 can include contouring, permanentlyreshaping or otherwise dimensioning the cover 186 according or withrespect to an external surface contour or profile of the main body 168of the component 160, such as the external surface contour CC of theairfoil section 62 of FIGS. 2-3. Various techniques can be utilized tocontour the cover 186, including hot forming and machining. The cover186 can be contoured with respect to a stagger angle of the respectiveairfoil that is twisted to define the predefined contour, including anyof the stagger angles disclosed herein, as illustrated by airfoil 61 ofFIGS. 2-3.

Forming the common cover 186 occurs such that a perimeter 186P of thecommon cover 186 is dimensioned to mate with the perimeter 78P of therecessed region 78 (shown in dashed lines in FIG. 8 for illustrativepurposes). In an installed position, the external surface contour of themain body 168 and external surfaces of the cover skins 170 can cooperateto define a pressure side or a suction side of an airfoil, asillustrated by cover skins 70 of FIGS. 2 and 3.

A plurality of cover skins 170 are formed from the common cover 186 atstep 184D. Step 184D includes dividing or segmenting the common cover186 (e.g., along the dashed lines illustrated in FIG. 8) to form thecover skins 170. The resultant cover skins 170 mounted to the main body168 are separate and distinct components. Various techniques can beutilized to form the cover skins 170 from the common cover 186, such aslaser cutting or another machining technique.

Step 184D can occur subsequent to contouring or otherwise forming thecommon cover 186 at step 184C, which can reduce manufacturing complexityin forming the cover skins 170 according to an external surface profileof the component mounting the cover skins 170. The peripheral cover skin172 comprises the perimeter 186P of the common cover 186 such that theperipheral cover skin 172 surrounds the localized cover skins 174 in theinstalled position, as illustrated by the cover skins 72, 74 of FIG. 2.

Step 184D can include forming pedestal cover skins containing one ormore recesses within the internal surfaces of a thicker-than-normalcommon cover 186 and/or cover skin 170. As illustrated in FIG. 8A, coverskin 170′ includes external surfaces 185′ and internal surfaces 187′opposed to the external surfaces 185′. The external surfaces 185′ candefine an external surface contour of the cover skin 170′, and internalsurfaces 187′ can bound a cavity or pocket 182′. The cover skin 170′ canbe chemically milled or otherwise machined to form a recess 193′. Thecover skin 170′ defines a first width W1 along a perimeter of the coverskin 170′ and defines a second width W2 along the recess 193′. Therecess 193′ can have a radiused transition from first width W1 to asecond width W2 as in FIG. 8A or can have a generally arcuate, concaveprofile as illustrated by recess 193″ of FIG. 8B such that first widthW1″ is greater than second width W2″ at a valley of the recess193′/193″. Incorporation of a pedestal cover skin can reduce the stressconcentration at the juncture of support rib 180′/180″ (shown in dashedlines in FIGS. 8A-8B for illustrative purposes) and cover skin170′/170″, which can result in improved fatigue life.

Referring to FIGS. 7 and 9-10, a perimeter of the cover skins 170 can bemachined or otherwise re-dimensioned subsequent to dividing the commoncover 186 according to a predefined geometry of the surface features ofthe component 160, such as the support ribs 80 and perimeter 78P of therecessed region 78 of the airfoil 61 of FIGS. 2-4.

Each of the localized cover skins 174 (one shown in FIG. 9 forillustrative purposes) and the peripheral cover skin 172 (FIG. 10) canbe mounted on a respective support structure 188, 189. Each supportstructure 188, 189 can be a vacuum chuck suctioning system having avacuum chuck contoured to a surface profile of the respective cover skin170. A machining assembly 190, 191 (shown in dashed lines forillustrative purposes) including a respective controller CONT and one ormore cutting tools CT can be utilized to resize or machine a perimeterof the respective cover skins 172, 174. Each cutting tool CT can be anendmill that is operable to mill the respective cover skin 172, 174 withrespect to a predefined geometry, for example. The cover skins 172, 174are re-dimensioned with respect to a width of the support ribs 180 andperimeter 178P of the recessed region 178 utilizing other techniques,such as laser cutting in an argon gas environment which can reduce alikelihood of surface contamination. Surfaces of the component 160 canbe cleaned prior to positioning the cover skins 170.

In the illustrative examples of FIGS. 9-10, a concave surface of thecover skins 170 mounted to the support structure 188, 189. In otherexamples, a convex surface of cover skin 170′ is mounted to supportstructure 188′/189′ as illustrated by FIG. 9A.

Referring to FIGS. 7 and 11, at step 184E the cover skins 170 arepositioned relative to the main body 168 including moving cover skins170′ (shown in dashed lines for illustrative purposes) in direction D1and into abutment with each adjacent rib 180 to enclose respective onesof the pockets 182. The cover skins 170 can be positioned such that theperipheral cover skin 172 surrounds the localized cover skins 174, asillustrated by cover skins 72, 74 of FIG. 2.

Each support rib 180 includes a neck portion 180A extending from a wallof the main body 168 and a pedestal portion 180B. The pedestal portion180B has a pair of shelves 180C that are dimensioned to support anopposed pair of the cover skins 170. In examples, the pedestal portion180B has a width of about 0.06-0.09 inches. Each rib 180 can include araised protrusion 180D extending outwardly from the pedestal portion180B to define a terminal portion of the rib 180. The pedestal portion180B can reduce stress concentrations at a junction between the rib 180and the respective cover skin 170.

The raised protrusion 180D is dimensioned to extend between, and spaceapart the cover skins 170. The raised protrusion 180D can be dimensionedto abut against the cover skins 170 in an installed position. Inexamples, the raised protrusion 180D has a width of approximately 0.025inches. In the illustrative example of FIG. 11, the raised protrusion180D is integral with the pedestal portion 180B. In other examples, theraised protrusion 180D is a separate and distinct component mechanicallyattached to the pedestal portion 180B of the respective rib 180.

In the illustrative example of FIG. 11, the raised protrusion 180D isdimensioned to extend outwardly of external surfaces of the cover skins170 subsequent to positioning the cover skins 170 to cover therespective pockets 182. The raised protrusion 180 can have a generallyrounded or tapered cross-sectional profile that serves to guide thecover skins 170′ toward the shelves 180C, which can reduce complexity inpositioning the cover skins 170 relative to the main body 168. In otherexamples, the raised protrusion 180D is omitted. Other geometries of theraised protrusion 180 can be utilized, such as raised protrusion 180D′having chamfered surfaces as illustrated in FIG. 11A and a generallyrectangular geometry as illustrated by raised protrusion 180D″ in FIG.11B.

Referring to FIGS. 7 and 12, at step 184F surfaces of each cover skin170 are mechanically attached to surfaces of the main body 168 along therespective rib 180 subsequent to positioning the cover skins 170 at step184E. Any of the techniques disclosed herein can be utilized tomechanically attach the cover skins 170 to the main body 168, includingwelding the cover skins 170 along the respective ribs 180 with a weldingsystem 192. The main body 168 can be mounted in a welding fixture. Thecover skins 170 are positioned relative to the main body 168 and heldagainst the main body 168 such that the ribs 180 directly abut againstthe cover skins 170 adjacent to the weld lines.

The cover skins 170 can be welded to the main body 168 along each raisedprotrusion 180D (FIG. 11), which is consumed during the welding suchthat the resulting weld 197 is slightly below or substantially flushwith the external surfaces of the adjacent cover skins 170, asillustrated by FIG. 12. Weld beam WB and respective edges of pedestalportion 180B′, raised protrusion 180D′, and cover skins 170′incorporated into the weld 197 are shown in dashed lines in FIG. 12 forillustrative purposes. The raised protrusion 180D provides integralfiller material to supplement weld metal drop-through that may occurduring formation of internal fillets 195 on either side of the rib 180.Utilizing the raised protrusion 180D to provide filler material mayserve to reduce a thickness of the adjacent cover skins 170, which mayotherwise be a relative greater thickness for underfill. A reduction inthickness may reduce material utilization and cost in fabricating thecomponent 160. The raised protrusion 180D can serve as a trackingfeature during welding, can reduce a depth of a surface depression inexternal surfaces of the component 160 adjacent the weld 197, and canreduce a need for attaching the cover skins 170 to the ribs 180 or otherportions of the main body 168 utilizing a blind weld technique.

Step 184F can include welding a perimeter of the peripheral cover skin172 to the main body 168 along a perimeter 178P of the recessed region178, as illustrated by FIG. 13. A width W3 of shelf 173 can be less thanor equal to a thickness or width W4 of cover skin 172, as illustrated inFIG. 13, which can reduce a size of the resultant weld. In examples, theperimeter of the cover skin 172 is welded to the perimeter 178P of therecessed region 178 prior to welding the localized cover skins 174,which can reduce overall distortion of the component 160. A stressrelief or creep form operation can be performed at step 184G to relievestresses in the component 160 caused by welding the cover skins 172, 174and main body 168. One or more finishing operations can be performed atstep 184H, including machining external surfaces of the component 160according to a predefined surface contour.

FIG. 14 illustrates a component 260 according to another example. In theillustrative example of FIG. 14, the component 260 is an airfoil 261including a main (or airfoil) body 268 having a plurality of supportribs 280 that bound or otherwise define a plurality of cavities orpockets 282 within a perimeter 278P of a recessed region 278. The ribs280 are arranged to define seven separate and distinct pockets 282(illustrated at 282-1 to 282-7).

A first set of the pockets 282-1 to 282-3 are dimensioned to have amajor component that extends in a chordwise direction X. Pockets 282-1to 282-3 are arranged to be generally parallel to each other. A secondset of the pockets 282-5 to 282-7 are dimensioned to have a majorcomponent that extends in a spanwise or radial direction R. Pockets282-5 to 282-7 are arranged to be generally parallel to each other andare generally traverse to the pockets 282-1 to 282-3.

Pocket 282-4 includes a first segment 282-4A and a second segment 282-4Bextending transversely from an end portion of the first segment 282-4A.The first segment 282-4A is dimensioned to have a major component thatextends in the radial direction R. The second segment 282-4B isdimensioned to have a major component that extends in the chordwisedirection X such that pocket 282-4 spaces apart the first set of pockets282-1 to 282-3 from the second set of pockets 282-5 to 282-7. Anotherpocket 282-8 follows or is otherwise defined by the perimeter 278P ofthe recessed region 278. Stiffening ribs 281 extend along a floor of thepocket 282-8 to provide rigidity to the main body 268. Pockets 282-1 to282-7 are free of any stiffening ribs. Each pocket 282 can be enclosedby a respective cover skin utilizing any of the techniques disclosedherein.

FIG. 15 illustrates a component 360 according to yet another example. Inthe illustrative example of FIG. 15, the component 360 is an airfoil361. A main (or airfoil) body 368 of component 360 defines a pluralityof cavities or pockets 382 (indicated at 382-1 to 382-4). Pockets 382-1,382-2 and 382-4 are dimensioned to have a major component that extendsin a spanwise or radial direction R. Pocket 382-3 includes a firstsegment 382-3A and one or more branched (second) segments 382-3B. Thebranched segments 382-3B extend outwardly at spaced intervals along alength of the first segment 382-3A. The first segment 382-3A can bedimensioned to have a major component that extends in the radialdirection R, and each branched segment 382-3B can be dimensioned to havea major component that extends in the chordwise direction X, forexample.

FIG. 16 illustrates a component 460 according to another example. In theillustrative example of FIG. 16, the component 460 is an airfoil 461.Main body 468 includes a plurality of support ribs 480 that define aplurality (or first set) of cavities or pockets 482 and a plurality (orsecond set) of stiffening ribs 481. The pockets 482 can have asubstantially similar profile as the pockets 382 of FIG. 15.

At least some of the stiffening ribs 481 extend between opposed walls ofa respective one of the pockets 482 encircled by the support ribs 480.At least some of the stiffening ribs 481 can be substantially alignedalong a common axis CA, as illustrated by stiffening ribs 481-1 to481-4. At least some of the stiffening ribs 481 can be joined at acommon node 494, as illustrated by stiffening ribs 481-2, 481-3, 481-5and 481-6.

Utilizing the techniques disclosed herein, the cover skins can beattached to the main body without the need for utilizing blind weldtechniques. The raised protrusion pedestal support rib can reduceoperating stresses at the weld joint, incorporate a consumable weldingtracking feature that provides additional weld filler to improve fillingof an external surface depression of the cover skin that may be createdby formation of internal fillets during welding, and can reducecomplexity in fabricating cover skins from a common workpiece or coverto mate with a geometry of a three-dimensionally twisted airfoil.

It should be understood that relative positional terms such as“forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like arewith reference to the normal operational attitude of the vehicle andshould not be considered otherwise limiting.

Although the different examples have the specific components shown inthe illustrations, embodiments of this disclosure are not limited tothose particular combinations. It is possible to use some of thecomponents or features from one of the examples in combination withfeatures or components from another one of the examples.

Although particular step sequences are shown, described, and claimed, itshould be understood that steps may be performed in any order, separatedor combined unless otherwise indicated and will still benefit from thepresent disclosure.

The foregoing description is exemplary rather than defined by thelimitations within. Various non-limiting embodiments are disclosedherein, however, one of ordinary skill in the art would recognize thatvarious modifications and variations in light of the above teachingswill fall within the scope of the appended claims. It is therefore to beunderstood that within the scope of the appended claims, the disclosuremay be practiced other than as specifically described. For that reasonthe appended claims should be studied to determine true scope andcontent.

What is claimed is:
 1. A method of forming a gas turbine enginecomponent comprising: forming a recessed region in a sidewall of a mainbody; dividing the recessed region into a plurality of pockets betweenone or more ribs such that the plurality of pockets are surrounded by aperimeter of the recessed region; contouring a cover according to anexternal surface contour of the main body such that a perimeter of thecover is dimensioned to mate with the perimeter of the recessed region;dividing the cover to form a plurality of cover skins subsequent to thecontouring step; positioning the plurality of cover skins to encloserespective ones of the plurality of pockets; and welding the pluralityof cover skins to the main body along the one or more ribs subsequent tothe positioning step; wherein the plurality of cover skins includes aperipheral cover skin and one or more localized cover skins, theperipheral cover skin comprising a perimeter of the cover such that theperipheral cover skin surrounds the one or more localized cover skinssubsequent to the positioning step; wherein the one or more ribs includea first set of ribs and a second set of ribs, each one of the second setof ribs extending from at least one of the first set of ribs, and thepositioning step includes situating one or more of the plurality ofcover skins over the second set of ribs; and wherein the second set ofribs are spaced apart from the plurality of cover skins.
 2. The methodas recited in claim 1, wherein the welding step includes welding theperipheral cover skin to the main body along the perimeter of therecessed region.
 3. The method as recited in claim 1, wherein each oneof the one or more ribs includes a raised protrusion that extendsoutwardly from a pedestal portion, the pedestal portion dimensioned tosupport an opposed pair of the plurality of cover skins, and the raisedprotrusion is dimensioned to abut against the opposed pair.
 4. Themethod as recited in claim 3, wherein the raised protrusion isdimensioned to extend outwardly from external surfaces of the opposedpair subsequent to the positioning step, and the raised protrusion isconsumed during the welding step.
 5. The method as recited in claim 1,wherein the external surface contour of the main body and externalsurfaces of the plurality of cover skins cooperate to define a pressureside or a suction side of an airfoil.
 6. A method of forming a gasturbine engine component comprising: forming a recessed region in asidewall of a main body; dividing the recessed region into a pluralityof pockets between one or more ribs such that the plurality of pocketsare surrounded by a perimeter of the recessed region; contouring a coveraccording to an external surface contour of the main body such that aperimeter of the cover is dimensioned to mate with the perimeter of therecessed region; dividing the cover to form a plurality of cover skinssubsequent to the contouring step; positioning the plurality of coverskins to enclose respective ones of the plurality of pockets; andwelding the plurality of cover skins to the main body along the one ormore ribs subsequent to the positioning step; wherein the step ofdividing the cover occurs such that the plurality of cover skinsincludes a peripheral cover skin and a plurality of localized coverskins, the peripheral cover skin comprising a perimeter of the cover;wherein the positioning step includes positioning the peripheral coverskin to surround the plurality of localized cover skins; wherein thewelding step includes welding the peripheral cover skin to the main bodyalong the perimeter of the recessed region; and wherein the positioningstep occurs such that the localized cover skins are spaced apart fromeach other and from the perimeter of the recessed region.
 7. The methodas recited in claim 6, wherein: the step of dividing the recessed regionincludes forming a second set of ribs in the recessed region such thateach rib of the second set of ribs extends from a respective rib of thefirst set of ribs; and wherein the positioning step includes situatingone or more of the plurality of cover skins over the second set of ribssuch that the second set of ribs are spaced apart from the plurality ofcover skins.
 8. The method as recited in claim 6, wherein the step ofdividing the cover occurs such that the plurality of localized coverskins exclude the perimeter of the cover.
 9. The method as recited inclaim 8, further comprising: forming the main body; and whereinpositioning step occurs such that the external surface contour of themain body and external surfaces of the plurality of cover skinscooperate to define a pressure side or a suction side of an airfoil. 10.The method as recited in claim 9, wherein the airfoil is a fan blade.11. The method as recited in claim 6, wherein: the step of dividing thecovers occurs such that the plurality of localized cover skins includesa first cover skin having a plurality of branch segments extending froman elongated body; the step of dividing the recessed region includesforming a first rib of the one or more ribs; and the positioning stepoccurs such that the first rib follows a perimeter of the first coverskin along each of the plurality of branch segments and along theelongated body.
 12. The method as recited in claim 1, furthercomprising: forming the main body to establish a portion of an airfoil;wherein the airfoil includes an airfoil body extending between leadingand trailing edges in a chordwise direction and extending from a rootsection in a spanwise direction, the airfoil body comprises the mainbody, and the airfoil body defines pressure and suction sides separatedin a thickness direction; and wherein the positioning step occurs suchthat the external surface contour of the main body and external surfacesof the plurality of cover skins cooperate to define the pressure side orthe suction side of the airfoil.
 13. The method as recited in claim 12,wherein the airfoil is a fan blade.